The present invention relates generally to gas turbine engines, and, more specifically, to turbine airfoil cooling.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through several turbine stages. A high pressure turbine (HPT) includes first stage turbine rotor blades extending outwardly from a supporting rotor disk which is rotated by the gases for powering the compressor. A low pressure turbine (LPT) follows the HPT and includes corresponding rotor blades which extract additional energy from the gases for performing useful work such as powering an output drive shaft. In one example, the shaft may be connected to a transmission for powering a military vehicle such as a battle tank.
Since the first stage turbine rotor blades are subject to the hottest combustion gas temperatures, they are cooled using a portion of the pressurized air bled from the compressor. However, any air bled from the compressor correspondingly decreases the overall efficiency of the engine, and therefore should be minimized.
The prior art contains a multitude of patents including various configurations for cooling turbine airfoils found in rotor blades or stator nozzle vanes. Various forms of cooling channels are known and include multi-pass serpentine cooling circuits, dedicated cooling channels for the leading edge or trailing edge of the airfoil, turbulators and pins for enhancing heat transfer by convection cooling, impingement cooling, apertures, and various forms of film cooling holes extending through the pressure and suction sidewalls of the airfoil.
The prior art is replete with different configurations for turbine airfoil cooling in view of the hostile operating environment in a gas turbine engine, and the substantial variation in heat loads from the combustion gases over the pressure and suction sides of the airfoil between the leading and trailing edges and root to tip thereof.
It is desired to maximize the cooling ability of the cooling air, while minimizing the amount of such cooling air diverted from the combustion process. Yet, sufficient air under sufficient pressure must be provided to the airfoils for driving the cooling air therethrough with sufficient pressure while maintaining sufficient backflow margin to prevent ingestion of the combustion gases through the various discharge holes in the airfoils. And, it is common to use the same cooling air for multiple cooling functions in a single turbine airfoil, which additionally increases the complexity of the design since the various cooling functions are then interrelated, with the upstream cooling features affecting the downstream cooling features as the cooling air absorbs heat along its flowpath.
A particularly difficult region of the turbine airfoil to cool is its leading edge along which the hot combustion gases first impinge the airfoil. The leading edge has an arcuate curvature which correspondingly creates more surface area on the external surface of the airfoil than its internal surface directly behind the leading edge in the first or leading edge flow channel located thereat. The leading edge flow channel may have smooth surfaces with impingement cooling thereof through a row of impingement holes in a forward bridge joining the pressure and suction sidewalls.
The spent impingement air is then typically discharged from the leading edge channel through multiple rows of film cooling holes typically arranged in a showerbead along the leading edge for providing external film cooling of the airfoil. Corresponding rows of gill holes may also be used downstream from the leading edge for additionally discharging the spent impingement air from the leading edge channel.
The leading edge channel may be otherwise configured with various forms of turbulators therein which protrude into the flow channel for tripping the cooling air channeled radially outwardly or inwardly depending upon the design.
Furthermore, stationary nozzle vanes may be cooled by channeling compressor bleed air either radially outwardly or inwardly therethrough. And, first stage turbine nozzles typically include impingement baffles suspended therein in yet another configuration for providing enhanced cooling thereof.
Correspondingly, turbine rotor blades receive their cooling air from the radially inner roots of the blades which are mounted around the perimeter of the rotor disk. Since the blades rotate during operation they are subject to substantial centrifugal forces which also affect performance of the cooling air being channeled through the blade airfoils.
Accordingly, it is desired to provide a turbine airfoil having improved internal cooling behind the leading edge thereof.